Compositions and methods of attachment of thick environmental barrier coatings on CMC components

ABSTRACT

A coating system on a CMC substrate is provided, along with methods of its tape deposition onto a substrate. The coating system can include a bond coat on a surface of the CMC substrate; a first rare earth silicate coating on the bond coat; a first sacrificial coating of a first reinforced rare earth silicate matrix on the at least one rare earth silicate layer; a second rare earth silicate coating on the sacrificial coating; a second sacrificial coating of a second reinforced rare earth silicate matrix on the second rare earth silicate coating; a third rare earth silicate coating on the second sacrificial coating; and an outer layer on the third rare earth silicate coating. The first sacrificial coating and the second sacrificial coating have, independently, a thickness of about 4 mils to about 40 mils.

FIELD OF THE INVENTION

The present invention relates generally to gas turbine engines turbines.More specifically, embodiments of the invention generally relate tothick environmental barrier coatings on CMC components, such as CMCblade tips.

BACKGROUND OF THE INVENTION

The turbine section of a gas turbine engine contains a rotor shaft andone or more turbine stages, each having a turbine disk (or rotor)mounted or otherwise carried by the shaft and turbine blades mounted toand radially extending from the periphery of the disk. A turbineassembly typically generates rotating shaft power by expanding hotcompressed gas produced by combustion of a fuel. Gas turbine buckets orblades generally have an airfoil shape designed to convert the thermaland kinetic energy of the flow path gases into mechanical rotation ofthe rotor.

Turbine performance and efficiency may be enhanced by reducing the spacebetween the tip of the rotating blade and the stationary shroud to limitthe flow of air over or around the top of the blade that would otherwisebypass the blade. For example, a blade may be configured so that its tipfits close to the shroud during engine operation. Thus, generating andmaintaining an efficient tip clearance is particularly desired forefficiency purposes.

Although turbine blades may be made of a number of superalloys (e.g.,nickel-based superalloys), ceramic matrix composites (CMCs)) are anattractive alternative to nickel-based superalloys for turbineapplications because of their high temperature capability and lightweight. However, CMC components must be protected with an environmentalbarrier coating (EBC) in turbine engine environments to avoid severeoxidation and recession in the presence of high temperature steam.

Thus, in certain components, regions of the EBC may be susceptible towear due to rub events with adjacent components. For example, for theCMC blade, the EBC at the blade tip is susceptible to rub against metalshroud components. If the EBC coating wears away, the CMC blade is thenopen to recessive attack from high temperature steam that will open upthe clearance between the CMC blade tip and the metal shroud, therebyreducing the efficiency of the engine.

Thus, it is desirable in the art to provide materials and methods forreducing EBC wear on a CMC blade tip caused by a rub event duringoperation of a turbine.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

A coating system on a CMC substrate is generally provided, along withmethods of its tape deposition onto a substrate. In one embodiment, thecoating system includes a first rare earth silicate coating on thesubstrate (e.g., with a bond coat positioned therebetween); a firstsacrificial coating of a first reinforced rare earth silicate matrix onthe at least one rare earth silicate layer; a second rare earth silicatecoating on the sacrificial coating; a second sacrificial coating of asecond reinforced rare earth silicate matrix on the second rare earthsilicate coating; a third rare earth silicate coating on the secondsacrificial coating; and an outer layer on the third rare earth silicatecoating. The first rare earth silicate coating, the second rare earthsilicate coating, and the third rare earth silicate coating include,independently, at least one rare earth silicate layer. The firstsacrificial coating and the second sacrificial coating have,independently, a thickness of about 4 mils to about 40 mils.

A blade (e.g., a turbine blade) is also generally provided that includesan airfoil having a CMC substrate and defining a blade tip, with atleast the blade tip of the airfoil including such a coating system.

In one embodiment, the method of tape deposition of a sacrificialcoating on a CMC substrate includes applying a first matrix materialonto a surface of a first film; drying the first matrix material toremove the solvent forming a first tape having a film side and a matrixside; applying the first tape onto the CMC substrate with the matrixside facing the CMC substrate; applying a second matrix material onto asurface of a second film applying the second tape onto the CMC substratewith the matrix side facing the CMC substrate; and sintering the firsttape and the second tape to bond the first matrix material to the CMCsubstrate via a first bonding layer and to bond the second matrixmaterial to the first matrix material via a second bonding layer.Generally, the first matrix material and the second matrix materialincludes, independently, a mixture of a rare earth silicate powder, asintering aid, and a solvent.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter which is regarded as the invention is particularlypointed out and distinctly claimed in the concluding part of thespecification. The invention, however, may be best understood byreference to the following description taken in conjunction with theaccompanying drawing figures in which:

FIG. 1 is a perspective view schematically representing an exemplaryturbine blade of a type formed of CMC materials;

FIG. 2 shows an exemplary coating system positioned on a blade tip of aturbine blade; and

FIG. 3 shows a cross-sectional view of the exemplary coating system ofFIG. 2 at the blade tip.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

Reference now will be made in detail to embodiments of the invention,one or more examples of which are illustrated in the drawings. Eachexample is provided by way of explanation of the invention, notlimitation of the invention. In fact, it will be apparent to thoseskilled in the art that various modifications and variations can be madein the present invention without departing from the scope or spirit ofthe invention. For instance, features illustrated or described as partof one embodiment can be used with another embodiment to yield a stillfurther embodiment. Thus, it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

In the present disclosure, when a layer is being described as “on” or“over” another layer or substrate, it is to be understood that thelayers can either be directly contacting each other or have anotherlayer or feature between the layers, unless expressly stated to thecontrary. Thus, these terms are simply describing the relative positionof the layers to each other and do not necessarily mean “on top of”since the relative position above or below depends upon the orientationof the device to the viewer.

Chemical elements are discussed in the present disclosure using theircommon chemical abbreviation, such as commonly found on a periodic tableof elements. For example, hydrogen is represented by its common chemicalabbreviation H; helium is represented by its common chemicalabbreviation He; and so forth. “Ln” refers to the rare earth elements ofscandium (Sc), yttrium (Y), lanthanum (La), cerium (Ce), praseodymium(Pr), neodymium (Nd), promethium (Pm), samarium (Sm), europium (Eu),gadolinium (Gd), terbium (Tb), dysprosium (Dy), holmium (Ho), erbium(Er), thulium (Tm), ytterbium (Yb), lutetium (Lu), or mixtures thereof.In particular embodiments, Ln is selected from the group consisting ofneodymium, gadolinium, erbium, yttrium, and mixtures thereof.

A coating for a CMC blade tip is generally provided herein, along withits methods of formation. The coating for the CMC blade tip isrelatively thick, dense, and mechanically resistant to spall and rub inturbine engine environments. The coating is deposited via attaching atleast two tapes filled with ceramic particles, sintering aids, organicbinders, and plasticizers.

The thick, tape-deposited sacrificial coating is generally provided incombination with a plurality of other, thinner layers to form an EBC ona CMC substrate. When applied to a blade tip, the sacrificial coatingprovides thickness that can rub away upon contact of the blade tip witha shroud. Thus, the sacrifice of this sacrificial coating during rubevents serves to protect the underlying layers of the EBC, such as arelatively thin underlayer of bond coat that in turn protects the CMCfrom oxidation and/or a relatively thin underlayer of rare earthdisilicate that in turn protects the CMC from high temperature steampenetration. It should also be noted that the sacrificial coatingitself, may also provide some protection against high temperature steampenetration.

In general, this overall coating system can be described as follows,from the CMC surface outward: a bond coat; one or more dense rare earthsilicate layer(s); a first thick layer (at least about 8 mils andpreferably at least about 15 mils) of rare earth disilicate matrix mixedwith a discontinuous phase of barium strontium aluminosilicate (BSAS) orsilicon metal particles (referred to herein as a “reinforced rare earthdisilicate matrix”); one or more dense rare earth silicate layer(s); asecond thick layer (at least about 8 mils and preferably at least about15 mils) of a reinforced rare earth disilicate matrix, and the option ofone or more rare earth silicate outer layer(s). Each of these layers isdescribed in greater detail below with respect to particular embodimentsprovided herein.

FIG. 1 shows an exemplary turbine blade 10 of a gas turbine engine. Theblade 10 is generally represented as being adapted for mounting to adisk or rotor (not shown) within the turbine section of an aircraft gasturbine engine. For this reason, the blade 10 is represented asincluding a dovetail 12 for anchoring the blade 10 to a turbine disk byinterlocking with a complementary dovetail slot formed in thecircumference of the disk. As represented in FIG. 1, the interlockingfeatures comprise protrusions referred to as tangs 14 that engagerecesses defined by the dovetail slot. The blade 10 is further shown ashaving a platform 16 that separates an airfoil 18 from a shank 15 onwhich the dovetail 12 is defined.

The blade 10 includes a blade tip 19 disposed opposite the platform 16.As such, the blade tip 19 generally defines the radially outermostportion of the blade 10 and, thus, may be configured to be positionedadjacent to a stationary shroud (not shown) of the gas turbine engine.As stated above, during use, the blade tip 19 may contact the shroud,causing a rub event between the blade tip 19 and the shroud.

In one particular embodiment, the blade tip 19 may be further equippedwith a blade tip shroud (not shown) which, in combination with tipshrouds of adjacent blades within the same stage, defines a band aroundthe blades that is capable of reducing blade vibrations and improvingairflow characteristics. By incorporating a seal tooth, blade tipshrouds are further capable of increasing the efficiency of the turbineby reducing combustion gas leakage between the blade tips and a shroudsurrounding the blade tips.

Because they are directly subjected to hot combustion gases duringoperation of the engine, the airfoil 18, platform 16 and blade tip 19have very demanding material requirements. The platform 16 and blade tip19 are further critical regions of a turbine blade in that they createthe inner and outer flowpath surfaces for the hot gas path within theturbine section. In addition, the platform 16 creates a seal to preventmixing of the hot combustion gases with lower temperature gases to whichthe shank 20, its dovetail 12 and the turbine disk are exposed, and theblade tip 19 is subjected to creep due to high strain loads and wearinteractions between it and the shroud surrounding the blade tips 19.The dovetail 12 is also a critical region in that it is subjected towear and high loads resulting from its engagement with a dovetail slotand the high centrifugal loading generated by the blade 10.

Referring to FIGS. 2 and 3, a coating system 20 is shown forming a thickEBC 22 on a CMC substrate 24 that defines the blade tip 19. In theexemplary embodiment shown, a bond coat 26 is positioned on the surface25 of the CMC substrate 24. A first rare earth silicate coating 28 a ison the bond coat 26 and is formed from at least one rare earth silicatelayer. A first sacrificial coating 30 a of a reinforced rare earthdisilicate matrix is positioned on the at least one rare earth silicatelayer 28 a. The first sacrificial coating 30 a has a thickness of about4 mils to about 40 mils (e.g., about 8 mils to about 25 mils, such asabout 16 mils to about 24 mils). A second rare earth silicate coating 28b is on the first sacrificial coating 30 a and is formed from at leastone rare earth silicate layer. As such, a rare earth silicate coating(collectively 28 a, 28 b) surrounds the first sacrificial coating 30 aat the blade tip 19. A second sacrificial coating 30 b of a reinforcedrare earth disilicate matrix is positioned on the second rare earthsilicate layer 28 b. The second sacrificial coating 30 b has a thicknessof about 4 mils to about 40 mils (e.g., about 8 mils to about 25 mils,such as about 16 mils to about 24 mils). A third rare earth silicatecoating 28 c is on the second sacrificial coating 30 b and is formedfrom at least one rare earth silicate layer. As such, a rare earthsilicate coating (collectively 28 b, 28 c) surrounds the secondsacrificial coating 30 b at the blade tip 19. Finally, an outer layer 32is positioned on the third rare earth silicate coating 28 c. Each ofthese layers is discussed in greater detail below. Also, it should benoted that more than two sacrificial coatings could be present in thecoating system 20, if desired.

As stated, the bond coat 26 is positioned in the CMC substrate 24, andin most embodiments is in direct contact with the CMC surface 25. Thebond coating generally provides oxidation protection to the underlyingCMC material 24. In one particular embodiment, the bond coat 26 is asilicon bond coat.

The first rare earth silicate coating 28 a generally provideshermeticity against high temperature steam. In one embodiment, the firstrare earth silicate coating 28 a is formed from at least one layer of aslurry-deposited yttrium ytterbium disilicate (YbYDS) layer. Othersilicate layers can be present in the first rare earth silicate coating28 a in order to provide hermeticity against high temperature steam,such as YbDS, LuDS, TmDS, LuYDS, TmYDS, etc. (where Lu=Lutetium andTm=Thulium), although any rare earth disilicate can be utilized.

The first sacrificial coating 30 a of a reinforced rare earth silicatematrix is generally formed by tape-depositing at least oneBSAS-reinforced rare earth silicate layer to the desired thickness, suchas about 4 mils to about 40 mils (e.g., about 8 mils to about 25 mils,such as about 16 mils to about 24 mils). Multiple layers may be utilizedto form the first sacrificial coating 30 a of the desired thickness. Thefirst sacrificial coating 30 a generally provides thickness to the EBC22 that can be sacrificed in a rub event by the blade tip 19 withanother component in the engine (e.g., a vane). The rare earth silicatelayers described with respect to the first sacrificial coating 30 a maybe comprised of rare earth disilicates (e.g., YbYDS), rare earthmonosilicates, or mixtures thereof.

The first sacrificial coating 30 a is deposited via a thicktape-deposition and sintering process, since it is very difficult tobuild up a thick coating on the tip of a blade by a thermal spraytechnique (since edge effects lead to spallation) or by slurrydeposition processes (since it would require multiple applications andheat treatments to build appreciable thickness). According to the thicktape-deposition method, the tape is loaded with the matrix material,such as the matrix material similar to that currently used for slurrydeposition of rare-earth disilicates. In this embodiment, a mixture ofrare earth disilicate powder and sintering aids that promote coatingdensification at temperatures of about 2300° F. to about 2500° F.(compared to about 2800° F. in the absence of the sintering aids) isutilized. In this method, however, a plurality of coarse particles(e.g., BSAS particles, silicon particles, or a mixture thereof) are alsoincluded in the tape so that they are at a level of about 30% to about65% by volume of the ceramic material, with the balance being the finerare earth silicate powder and sintering aid. The coarse particles have,in one embodiment, an average particle size of about 5 microns to about100 microns. The coarse particle addition helps overcome the problem ofthe slurry process such that one obtains a thick, crack free layer afterheat treatment. The use of BSAS or silicon coarse particles,specifically, also helps keep the porosity in the layer low (on theorder of about 20% by volume or less, and in some embodiments, as littleas about 10% porosity or less). Other coarse particles, such as ZrO₂,can result in porosity levels above 20% by volume. The matrix materialalso contains organic binder (e.g., polyvinyl butyral) and plasticizer(e.g., dibutyl phthalate or dipropylene glycol dibenzoate) so that thetape is flexible and tacky for the attachment to the CMC blade tipsurface. The tape is formed from slurry that comprises all of theconstituents mentioned above, plus one or more solvents. The slurry canbe cast under a doctor blade with a gap set to a controlled thickness,onto a film (e.g., a polymeric film). The solvent is then removed bydrying, yielding the tape. In certain embodiments, the dryingtemperature is about 15° C. to about 50° C., and can be dried at roomtemperature (e.g., about 20° C. to about 25° C.). Drying can beaccomplished for any suitable duration (e.g., about 30 minutes to about50 hours). Thus, another advantage of the tape approach is that there isno drying process after the tape is attached to the blade tip thatresult in drying defects that alter the geometry of the thick tip.

The tapes can be transferred to the CMC substrate by any suitablemethod. For example, the tape can be transferred to the CMC substratethrough applying pressure in combination with the tack of the tape orthrough applying pressure in combination with an elevated temperature,to get the tape to flow a bit into the roughness of the blade tipsurface, and the tack of the tape. In either of these methods, theadditional application of a solvent to the tape surface can increase itstack. In one particular embodiment, these methods can be utilized withthe addition of slurry, such as rare earth disilicate and sintering aidsbut without the BSAS particles. The addition of the slurry tends tocreate a robust bond during sintering.

Multiple tape transfers can be performed, in particular embodiments, tobuild the resulting first sacrificial coating 30 a to the desiredthickness.

The second rare earth silicate coating 28 b also provides hermeticityagainst high temperature steam. In one embodiment, the second rare earthsilicate coating 28 b is formed from at least one layer of aslurry-deposited yttrium ytterbium disilicate (YbYDS) layer. Othersilicate layers can be present in the second rare earth silicate coating28 b, similar to those described above with respect to the first rareearth silicate coating 28 a in order to provide hermeticity against hightemperature steam.

In one particular embodiment, the first rare earth silicate coating 28 aand the second rare earth silicate coating 28 b are substantiallyidentical in terms of composition. Referring to FIG. 2, the first rareearth silicate coating 28 a and the second rare earth silicate coating28 b are extensions of the same rare earth silicate coating 28, but fortheir respective positioning to surround the sacrificial coating 30 atthe blade tip 19. As shown, the separation points 29 a serve to splitthe rare earth silicate coating 28 into the first rare earth silicatecoating 28 a and the second rare earth silicate coating 28 b positionedabout the first sacrificial coating 30 a. In this embodiment, the firstsacrificial coating 30 a is completely encased within the first rareearth silicate coating 28 a and the second rare earth silicate coating28 b in order to form a hermetic seal against high temperature steam.Additionally, the second rare earth silicate coating 28 b may provideadditional mechanical stability for the underlying first sacrificialcoating 30 a (e.g., formed from a BSAS-reinforced YbYDS layer).

Both the first rare earth silicate coating 28 a and the second rareearth silicate coating 28 b can be formed via slurry deposition. In oneembodiment, the first rare earth silicate coating 28 a is deposited,followed by tape-deposition of the first sacrificial coating 30 a in thelocation desired. Then, the second rare earth silicate coating 28 b canbe deposited (e.g., via slurry deposition) onto the first sacrificialcoating 30 a and the exposed first rare earth silicate coating 28 a.Where there is no first sacrificial coating 30 a present (e.g., on theleading edge, the blade surface, the trailing edge, etc.), the secondrare earth silicate coating 28 b is merged with the first rare earthsilicate coating 28 a in order to form a single layer of the rare earthsilicate coating 28.

As shown in FIGS. 2 and 3, the second sacrificial coating 30 b of areinforced rare earth silicate matrix is generally formed bytape-depositing at least one BSAS-reinforced rare earth silicate layerto the desired thickness, such as about 4 mils to about 40 mils (e.g.,about 8 mils to about 25 mils, such as about 16 mils to about 24 mils).Multiple layers may be utilized to form the second sacrificial coating30 b of the desired thickness. The second sacrificial coating 30 bgenerally provides thickness to the EBC 22 that can be sacrificed in arub event by the blade tip 19 with another component in the engine(e.g., a vane). The rare earth silicate layers described with respect tothe second sacrificial coating 30 b may be comprised of rare earthdisilicates (e.g., YbYDS), rare earth monosilicates, or mixturesthereof. The second sacrificial coating 30 b may be formed by any of themethods discussed above with respect to the first sacrificial coating 30a.

The third rare earth silicate coating 28 c provides hermeticity againsthigh temperature steam. In one embodiment, the third rare earth silicatecoating 28 c is formed from at least one layer of a slurry-depositedyttrium ytterbium disilicate (YbYDS) layer. Other silicate layers can bepresent in the third rare earth silicate coating 28 c, similar to thosedescribed above with respect to the first rare earth silicate coating 28a in order to provide hermeticity against high temperature steam.

In one particular embodiment, the second rare earth silicate coating 28b and the third rare earth silicate coating 28 c are substantiallyidentical in terms of composition. Referring to FIG. 2, the second rareearth silicate coating 28 b and the third rare earth silicate coating 28c are extensions of the same rare earth silicate coating 28, but fortheir respective positioning to surround the second sacrificial coating30 b at the blade tip 19. As shown, the separation points 29 b serve tosplit the rare earth silicate coating 28 into the second rare earthsilicate coating 28 b and the third rare earth silicate coating 28 cpositioned about the second sacrificial coating 30 b. In thisembodiment, the second sacrificial coating 30 b is completely encasedwithin the second rare earth silicate coating 28 b and the third rareearth silicate coating 28 c in order to form a hermetic seal againsthigh temperature steam. Additionally, the third rare earth silicatecoating 28 c may provide additional mechanical stability for theunderlying second sacrificial coating 30 b (e.g., formed from aBSAS-reinforced YbYDS layer).

Like both the first rare earth silicate coating 28 a and the second rareearth silicate coating 28 b, the third rare earth silicate coating 28 ccan be formed via slurry deposition. In one embodiment, the second rareearth silicate coating 28 b is deposited, followed by tape-deposition ofthe second sacrificial coating 30 b in the location desired. Then, thethird rare earth silicate coating 28 c can be deposited (e.g., viaslurry deposition) onto the second sacrificial coating 30 b and theexposed first second earth silicate coating 28 b. Where there is nosecond sacrificial coating 30 b present (e.g., on the leading edge, theblade surface, the trailing edge, etc.), the third rare earth silicatecoating 28 c is merged with the second rare earth silicate coating 28 bin order to form a single layer of the rare earth silicate coating 28.

Thus, in one embodiment, each of the first rare earth silicate coating28 a, the second rare earth silicate coating 28 b, and the third rareearth silicate coating 28 c are made with a substantially identicalcomposition.

Finally, an outer layer 32 is positioned on the third rare earthsilicate coating 28 c. In one embodiment, the outer layer 32 comprisesat least one slurry-deposited yttrium monosilicate (YMS) layer. Theouter layer 32 provides protection against steam recession and moltendust. Materials other than rare earth silicates can be utilized withinthe outer coating, such as rare earth hafnates, rare earth zirconates,rare earth gallates (e.g., monoclinic type, such as Ln₄Ga₂O₉), rareearth monotitanate (e.g., Ln₂TiO₅), rare earth cerate (e.g., Ln₂CeO₅),rare earth germinate (e.g., Ln₂GeO₅), or mixtures thereof. However, allof these materials have a relatively high coeffiecient of thermalexpansion (CTE) compared to rare earth silicate. Thus, rare earthmonosilicate is preferred. Hafnia, rare-earth stabilized hafnia, andrare-earth stabilized zirconia provide protection against steamrecession but not CMAS, and also have higher CTE than rare earthmonosilicate.

In addition to a thick coating on the blade tip 19, the EBC 20 can beused as an alternate method to obtain a thick EBC coating on othercomponents or areas of a CMC component (e.g., on a shroud, etc.).

While the invention has been described in terms of one or moreparticular embodiments, it is apparent that other forms could be adoptedby one skilled in the art. It is to be understood that the use of“comprising” in conjunction with the coating compositions describedherein specifically discloses and includes the embodiments wherein thecoating compositions “consist essentially of” the named components(i.e., contain the named components and no other components thatsignificantly adversely affect the basic and novel features disclosed),and embodiments wherein the coating compositions “consist of” the namedcomponents (i.e., contain only the named components except forcontaminants which are naturally and inevitably present in each of thenamed components).

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed:
 1. A coating system on a CMC blade having a blade tip,the coating system comprising: a first rare earth silicate coating onthe substrate, wherein the first rare earth silicate coating comprisesat least one rare earth silicate layer; a first sacrificial coating of afirst reinforced rare earth silicate matrix on the at least one rareearth silicate layer, wherein the first sacrificial coating has athickness of about 4 mils to about 40 mils, and wherein the firstreinforced rare earth silicate matrix comprises a rare earth silicatematrix with a discontinuous phase of barium strontium aluminosilicate orsilicon metal particles; a second rare earth silicate coating on thesacrificial coating, wherein the second rare earth silicate coatingcomprises at least one rare earth silicate layer, and wherein the firstrare earth silicate coating and the second rare earth silicate coatingsurround the first sacrificial coating at the blade tip so as to mergetogether; a second sacrificial coating of a second reinforced rare earthsilicate matrix on the second rare earth silicate coating, wherein thesecond sacrificial coating has a thickness of about 8 mils to about 25mils, and wherein the second reinforced rare earth silicate matrixcomprises a rare earth silicate matrix with a discontinuous phase ofbarium strontium aluminosilicate or silicon metal particles; a thirdrare earth silicate coating on the second sacrificial coating, whereinthe third rare earth silicate coating comprises at least one rare earthsilicate layer, and wherein the second rare earth silicate coating andthe third rare earth silicate coating surround the second sacrificialcoating at the blade tip so as to merge together; and an outer layer onthe third rare earth silicate coating.
 2. The coating system of claim 1,wherein the first sacrificial coating has a thickness of about 8 mils toabout 25 mils.
 3. The coating system of claim 1, wherein the firstsacrificial coating has a thickness of about 16 mils to about 24 mils.4. The coating system of claim 1, wherein the reinforced rare earthsilicate matrix of the first sacrificial coating and of the secondsacrificial coating comprises a rare earth silicate mixed with adiscontinuous phase of barium strontium aluminosilicate.
 5. The coatingsystem of claim 1, wherein the reinforced rare earth silicate matrix ofthe first sacrificial coating and of the second sacrificial coatingcomprises a rare earth silicate mixed with a discontinuous phase ofsilicon metal particles.
 6. The coating system of claim 1, wherein thereinforced rare earth silicate matrix of the first sacrificial coatingand of the second sacrificial coating comprises a rare earth disilicate.7. The coating system of claim 1, wherein the reinforced rare earthsilicate matrix of the first sacrificial coating and of the secondsacrificial coating comprises a rare earth monosilicate.
 8. The coatingsystem of claim 1, wherein the reinforced rare earth silicate matrix ofthe first sacrificial coating and of the second sacrificial coatingcomprises a rare earth disilicate, a rare earth monosilicate, or amixture of a rare earth monosilicate and a rare earth disilicate.
 9. Thecoating system of claim 1, further comprising: a bond coat on a surfaceof the CMC substrate between the CMC substrate and the first rare earthsilicate coating, wherein the bond coat is a silicon bond coat.
 10. Thecoating system of claim 1, wherein the first rare earth silicatecoating, the second rare earth silicate coating, and the third rareearth silicate coating have a substantially identical composition andare extensions of the same rare earth silicon coating but for theirrespective positioning to surround the first sacrificial coating betweenthe first rare earth silicate coating and the second rare earth silicatecoating and to surround the second sacrificial coating between thesecond rare earth silicate coating and the third rare earth silicatecoating.
 11. A blade, comprising: an airfoil comprising a CMC substrateand defining a blade tip, wherein a blade surface is defined along aleading edge and a trailing edge of the airfoil; wherein the blade tiphas the coating system of claim 1 thereon such that the firstsacrificial coating and the second sacrificial coating are on the bladetip while the first rare earth silicate coating, the second rare earthsilicate coating, and the third rare earth silicate coating form thesame rare earth silicon coating on blade surface.
 12. The blade as inclaim 11, wherein the blade is a turbine blade comprising: a shankdefining a dovetail; and a platform separating the airfoil from theshank.